Heat dispersing swivel rocket nozzle liner



Aug. 9, 1966 H. E. HELMs r-:TAL 3,265,314

HEAT DISPERSING SWIVEL ROCKET NOZZLE LINER Filed sept. 15. 196:5

INVENTORS Ward/a .Ve/1715,

7E/M25 d @me ATTORNEY United States Patent Office Patented August 9, 1966 3,265,314 HEAT DISPERSTNG SWllVEL ROCKET NOZZLE LINER Harold E. Helms, Indianapolis, Kenneth O. Johnson,

Camby, and .lames A. Pyne, Indianapolis, Ind., assignors to General Motors Corporation, Detroit, Mich., a corporation of Delaware Filed Sept. 13, 1963, Ser. No. 308,906 1 Claim. (Cl. 239-591) This invention relates to rocket nozzle-s using a new principle of heat sink structure and, more particularly, to a nozzle liner system which results in an effective heat sink for a longer period of time than results from the conventional heat sink design.

Several problems have arisen in present day rocket engine nozzles with respect to failure of the nozzle structure 'when subjected to extremely high temperatures. For the heat sink nozzle design, he-at is conducted radially outward, from the hotter heat transfer surface, into the cooler heat sink material. The nozzle heat transfer surface temperature increases very rapidly with time to a very high value :and then increases more slowly, gradually approaching the gas temperature, as the firing cycle progresses. The rest of the heat sink material lnehaves in a simil-ar fashion `but ywith a different time rate which depends upon the distance away `from ythe heat transfer surface, the type `of fhe-at sink material or materials and the-ir arrangement. This results in large ythermal stresses, localized cracking, and failure. As the lsurface temperature increases, the surface layers of the heat sink material become weak and disappear due to erosion and phase change. If the he'at sink were effective for a l-onger period of time, these phenomena would not occur until later in the tiring cycle, ideally, they would occur -just at the end of the tiring cycle. Thus, it has become ,apparent that a nozzle liner system is needed which will prolong the effectiveness of the heat sink, thereby reducing the possibility of erosion and failure of the nozzle.

Therefore, it is an object of this invention to provide a nozzle liner system for heat sink nozzles which will prolong the usefulness of the heat sink. This will produce lower temperatures in the wall at a `given time and thus reduce the detrimental effec-ts due to high temperature. These detrimental effects include but lare not limited to localize-d cracking, erosion and burn-out.

Other objects, ffeatures land `advantages of the invention will Ibecome readily apparent upon .reference to the succeeding detailed description and the drawings illustrating the preferred embodiment thereof; Within,

FIGURE 1 is a schematic View of the rocket en-gine showing the nozzle liner system in section; and

FIGURE 2 is `an enlarged view of the nozzle liner system of FIGURE fl.

In general, then, the invention relates to -a nozzle liner sys-tem lwhich provides la heat ilow path which is not purely radial and which produces beneficial effects. This heat flow path is provided by la combination of several liner components which will be described.

Mo-re particularly, FIGURE 1 shows a rocket engine xhaust nozzle 14. It is of the vectoring type and may rotate lwithin the nozzle inlet housing 16. This vectoring action is provided by having the nozzle |14 pivot on pin members 18, which are Iattached to the housing 16. The nozzle 14 is of the convergi-ng-diverging type having a converging inlet portion 20, la throat 40r venturi portion 22, and ya diverg'ing exit cone V24. 'Ilhe invention is applied herein to a vectoring nozzle; however, it is readily recognized that it is not llimited -to a vectoring type of nozzle. The nozzle liner system is shown in detail in FIGURE 2. The nose portion 26 of the nozzle 14 has a tungsten liner 28. This liner wraps around the nom and extends around the exterior of the main nozzle liner. The main body of the liner syste-m is composed of carbon rings 30 of normal vgraphite material. The surface of the converging portion 20 of the nozzle 14 has a pyrolytic graphi-te liner 32. The throat section `22 of the nozzle 14 is lined with another tungsten liner 34. This ltungsten liner 34 has a conical ange 36 which extends between the carbon rings l3l). The carbo-n rings 30 are backed up by the normal plastic substance 38. The remaining liner members of the nozzle will Ibe of structural material such as tantalum, titanium and zirconia. The zirconia forms a lthin layer 40 .between the carbon block 30 and the plastic substance 38. The titanium provides 'a hacking structure 42 lfor the plastic sub-stance '38 and the tantalum comprises the rbearing surface 44 upon which the nozzle structure slides when vectoring.

The pyrolytic Vgraphite liner 32 used in the converging portion I20 of the exhaust nozzle 14 has heat conduction properties which make it very useful in this portion. Pyrolytic graphite is known to have very high heat conducting properties along one axis while Ihaving very low heat conducting properties perpendicular to this axis. Therefore, with the proper orientation of :the `grain structure of the pyrolytic graphite the flow of heat can be restricted to a desired direction. The use, then, of ythis pyrolytic graphite liner 312 in combination with the tungsten liner 34 and the graphite rings 30 provides the line-r system previously mentioned. The converging portion 20 of the nozzle liner normally has been carbon rings with isotropic thermal properties |which conduct heat radially 'and axially equally 'wel-l. The thermal stress generated hy the thermal gradient in the graphite has caused cracking which at times was severe enough to result in the loss of chunk-s of material which was 'followed `by localized burnout and catastrophic failure. The pyrolytic .graphite liner 32 is chosen such that its good heat conducting direction is parallel to the tlow of ho-t gases and its poor heat conducting direction is normal to the flow of hot gases, thereby restricting the radial inward flow of heat in this area. Therefore, as the hot gas contacts the pyrolytic graphite liner 32 it increases in temperature but only a small portion of the heat ows through this liner 32 t-o the normal graphite 30 tbehind it. Thus, the orientation of the pyrolytic graphite 32 results in a heat shield which prolongs the usefulness of the heat sink 30. The tungsten insert 34 extends into the cool carbon rings 30 land heat is conducted down the tungsten insert and its .radial extension 36 :and dispersed in the car-bon rings V30. This reduces the rate of temperature increase of the nozzle Iheat sink and make-s the radial temperature distribution in the graphite rings more uniform, thus reducing the :thermal stresses in these rings.

Analysis 'ha-s shown that the surface temperature of the throat section 2.2v ha-s been reduced from 4800 F. to 4035 F. at the end of .sixty seconds when the subject nozzle system is used in place of a normal carbon nozzle liner. rThis reduction is very important when it is realized that the tungsten form-s an alloy with the carbon which melts at 4600 F. Since the throat section 22 is the most critical tarea rfor determining the resultant thrust of the rocket engine, it is highly desirable that this area remain intact as long as possible. Therefore, the reduction of the temperature of this area Iby the subject nozzle liner system insures a longer life for this port-ion, Iand thereby enables the rocket engine to produce maximum thrust for a correspondingly longer period of time.

In summary then, it is seen that the subject nozzle liner system provides an effective method of controlling the rate of |heat flow through the nozzle liner of a rocket engine and it further controls the heat flow such :that the nozzle liner will be less susceptible to thermal shock and associated cracking. The ultimate result of this is that the throat 4portion of the nozzle liner remains intact for a tlonger period of time at the same gastemperature or, for the lsame period lof time, at a higher gas temperature.

Although the subject invention has been illustra-ted with respect to a nozzle liner of an exhaust nozzle of a rocket engine, it will be clear to those skilled in the art to which the invention pertains that the lsubject invention will have imany applications other than that which is disclosed in which a heat lflow path controlling liner system is desired, and that many modieations and changes may be made thereto without departing from the scope of the invention.

We claim:

A rocket engine vectoring exhaust nozzle having a nose portion, a converging portion, a throat portion, 'and a heat flow path controlling liner syste-m comprising, in combination, a `carbon masts forming the larger portion of the nozzle structure, a tungsten liner at the nose port-ion of the nozzle, a pyrolytic graphite liner shielding the converging porti-on off the nozzle, said pyrolytlic graphite line-r portion 'having its one-dimensional heat conduction characteristics oriented to cause the heat to `flo-vv `along the nozzle surface, parallel to the flow of gases through the nozzle, thereby restricting radial heat flow, a tungsten liner expose-d at the throat portion of the nozzle having ya poirtion extending radially and toward the nose portion throughout the carb-on mass thereby reducing the possibili-ty of thermal shock and loeal failures.

References Cited by the Examiner UNITED STATES PATENTS `2,849,860 9/1958 Lowe 60e-35.6 3,106,061 10/1963 Eder 60-3555 3,115,746 12/1963 Hsia 60-35.6 3,137,132 6/1964 Turkat 60-35.6 3,156,091 11/1964 Kraus 6035.6 3,165,888 =1/1965 Keen 60-35.6

FOREIGN PATENTS 1,246,339 10/1960 France.

OTHER REFERENCES -Rooket Retractories bly Porter, Navord Report 4893, Notts. 1191, August 1955, pp. 9-12 relied on.

MARK NEWMAN, Primary Examiner.

CARLTON R. CROYLE, Examinez'. 

